The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. The gases flow through turbine stages which extract energy therefrom for powering the compressor and producing useful work such as powering a fan in an aircraft turbofan gas turbine engine.
Each turbine stage includes a stator nozzle having vanes which direct the combustion gases against a corresponding row of turbine blades extending radially outwardly from a supporting rotor disk. The vanes and blades include airfoils having generally concave pressure sidewalls and generally convex suction sidewalls extending axially between leading and trailing edges over which the combustion gases flow during operation.
The turbine blades are mounted to the rotor disk with corresponding dovetails which engage complementary dovetail slots formed in the perimeter of the disk. Each blade includes an inboard platform defining the radially inner boundary of the combustion gas flowpath, with the airfoil extending from a root thereat to a radially outboard tip. The blade tips are spaced closely adjacent to a surrounding stationary shroud for reducing leakage of the combustion gases in the gap therebetween during operation.
However, due to differential expansion and contraction between the blades and surrounding shroud during operation, the blade tips are subject to occasional tip rubs with the shroud.
In order to protect the blade tips, they are typically configured in the form of a squealer rib extension of the pressure and suction sidewalls which extends radially outwardly from a tip cap or floor that closes the radially outer end of the airfoil. The airfoil is hollow below the tip cap and includes various cooling channels or circuits therein for channeling air bled from the compressor for use as a coolant against the heating effect of the hot combustion gases.
In this configuration, the squealer ribs provide short extensions of the airfoil sidewalls for maintaining the aerodynamic profile thereof and provide minimum contact area with the shroud during tip rubs therewith. The underlying tip cap is therefore spaced further away from the shroud and is protected during tip rubs for maintaining the integrity of the airfoil, including the cooling channels therein.
During operation, the squealer ribs are directly subject to the hot combustion gases which flow thereover through the gap with the turbine shroud. They are therefore subject to heating on their three exposed sides, and are correspondingly difficult to cool. High temperature operation of the squealer ribs adversely affects the useful life thereof. The squealer ribs are cooled by conduction radially inwardly through the airfoil sidewalls with the heat being removed in the coolant channeled inside the airfoil. The airfoil may also include inclined tip holes disposed radially inwardly of the squealer ribs for forming a film cooling boundary of air typically along the pressure side of the airfoil for protecting the pressure side squealer rib portion.
Since the squealer ribs are disposed on both sides of the airfoil above the tip cap, they define therebetween an open tip cavity in which hot combustion gases may circulate to heat the inner sides of the squealer rib. The tip cap may include holes therethrough for discharging a portion of the coolant through the tip cavity, yet the squealer rib is still subject to heating on its three exposed sides.
Accordingly, it is desired to provide a turbine airfoil having improved tip cooling for increasing the useful life thereof or permitting operation with higher temperature combustion gases.